Turbine engines may be used to power various types of flight vehicles, including supersonic and hypersonic air and space vehicles and weaponry. Generally, gas turbine engines generate compressed air from a compressor section that enters a combustor section where an array of fuel nozzles injects a steady stream of fuel into the compressed air stream. The compressed air and fuel mixture is then ignited by one or more burners causing rapidly expanding air to flow through a turbine section, which in turn causes rotationally mounted turbine blades to rotate and generate energy to drive the compressor section, among others. The turbine section may have multiple stages to more efficiently extract energy from the airflow. The air exiting the turbine section is exhausted from the engine via an exhaust section, which may include an afterburner, thereby creating thrust. Supersonic and hypersonic flight requires that the turbine engine rapidly burn a considerable amount of fuel and rapidly displace a considerable amount of air, which in turn, leads to a significant amount of friction, and thus, heat generated by the turbine engine.
The output power of a turbine engine may be controlled by metering the fuel and air flows into the engine. To burn high flow volumes, modern gas turbine engines divide the fuel flow and burning into several (e.g., 6, 8 or 10) injection sites or combustion zones. The combustion zones are typically arranged in an array (e.g., a ring pattern), such that burned fuel in each combustion zone provides a flame front that effects a pressure change that drives the turbine blades. The pressure differential is dependent upon the temperature of the flame front. The higher the flame temperature, the greater the change in pressure, and thus the more power output from the turbine engine. The overall flame temperature is actually an average of the flame temperature at each burner or combustion zone. The temperature gradient profile of the several burners is defined by its “pattern factor,” which is typically defined as the difference between the peak and average combustor exit temperatures divided by the average exit temperature.
Ideally, the average flame temperature of all combustion zones should equal the flame temperature at the flame front so that the pattern factor is zero. However, practically, the average temperature is some valve less than the peak temperature, resulting in a positive pattern factor value. Should one or more combustion zones have a significantly relatively lower temperature than the others, the average flame front temperature can vary significantly from the peak temperature, thereby resulting in a high pattern factor, and inefficient operation of the turbine as well as the possible generation of pressure oscillations that may impart vibrations or other mechanical anomalies to the moving components (e.g., various vanes and blades) of the turbine engine.
High pattern factors, or temperature profile variations, may result from inconsistent fuel flow to the various injector nozzles of the combustion zones. Inconsistent fuel flow may result from even slight differences in the dimensioning or tolerances of the flow valves as well as from deterioration (e.g., coking) due to the contaminated and aggravated temperature environments in which the valves and nozzles are operated. Systems for actively controlling the turbine pattern factor may include electronic controls that use temperature feedback signals at the injector nozzles. Yet, such systems, for example those in large-scale power generating gas turbines, may introduce cost, weight, and failure points to the system, and may be insufficiently responsive or accurate to perform adequately at the high flow rates and pressures experienced in hypersonic flight applications to achieve the desired pattern factor control of the burner temperature profile.